The use of high strength fiber reinforced composite materials in the manufacture of aircraft and other lightweight structures has increased steadily since the introduction of such materials. Composite materials have a high strength-to-weight ratio and stiffness. These properties make composite materials attractive for use in the design of lightweight structures. Some of the drawbacks to using composite materials have been relatively high fabrication costs and difficulties in manufacturing defect-free parts. Generally, it has been difficult to produce parts formed of high strength composite materials at the same cost as comparable metal parts. It has also been difficult to fabricate composite parts in which tight dimensional tolerances are required on opposing surfaces or multiple surfaces of the formed composite part.
When it is necessary to fabricate composite parts having tight dimensional tolerances on multiple surfaces, it is generally necessary to use matched rigid tooling. In matched rigid tooling, two or more tools having forming surfaces that establish the final part dimensions are used. Once assembled, the forming surfaces of such rigid tooling define the final formed composite part's dimensions. Therefore, matched rigid tooling must be carefully fabricated to ensure that the tools fit together properly to ensure proper part dimensions. Inaccuracies in either the way the tools fit together or in the forming surfaces on any of the tools can result in defects in the formed composite part. Because of the tight machining tolerances required on matched rigid tooling, such tooling tends to be expensive to fabricate, thus increasing overall part fabrication costs.
The difficulties associated with fabricating high quality composite parts with rigid matched tooling is increased by variations in the quality and thickness of uncured composite materials. High-strength composite parts in the aerospace industry are generally fabricated using composite prepreg material. Such prepreg material consists of unidirectional fibers or cloths of graphite, fiberglass, Kevlar.RTM., etc., contained within a matrix material such as an epoxy, bismaleimide, or thermoplastic material. The thickness of individual layers of composite prepreg differs slightly between batches of prepreg and even within the same batch of prepreg. Therefore, the thickness of a workpiece formed of multiple layers of prepreg differs from workpiece to workpiece. Such differences in thickness from workpiece to workpiece can result in differences in the thickness of the final formed composite parts.
When matched rigid tooling is used, the forming surfaces of the tools generally cannot account for the differing thicknesses in the prepreg used. Therefore, if the prepreg used to form the composite part is slightly thicker than normal, portions of the formed composite part will have a lower percentage of resin than desirable. In extreme cases, the increased thickness of the composite prepreg can prevent the matched rigid tooling from fitting together properly, thus leading to tolerance problems on the surface of the formed composite part. Similarly, if the prepreg used is slightly thinner than normal, the use of the matched rigid tooling does not account for the thickness variations, thus resulting in a resin-rich formed composite part.
It would be desirable if a system of composite tooling was developed that could produce composite parts with tight tolerance control on multiple surfaces. At the same time, it would be advantageous if such a tooling method could eliminate the need for expensive matched rigid tooling.
As can be seen from the above discussion, there exists a need for a method of forming composite parts that can maintain dimensional tolerances, while reducing the disadvantages of the prior art. The present invention is directed toward fulfilling this need.